Material assembly for an inflatable aerodynamic braking device for spacecraft deceleration and the like

ABSTRACT

The invention is a material for use in an inflatable aerodynamic braking device for a spacecraft. In detail, the invention includes a plurality of high temperature resistant barrier outer layers. Preferably, the outer layer is made of Aluminoborosilicate fabric. At least one carbon cloth layer is disposed behind said outer layers. A plurality of thermal radiation barrier intermediate layers are disposed behind the at least one carbon cloth layer. The intermediate layers include a plurality of metal foil layers, at least one additional carbon cloth layer interspersed between the metal foil layers. A plurality of metal-coated polymeric layers are disposed behind the metal foil layers, preferably a polyimide fabric coated with Aluminum. The metal foil layers are Aluminum and Titanium. At least one inflatable gas barrier interior layer is disposed behind the intermediate layers, preferably made of a polyimide material. The outer, intermediate and interior layers are sewn together by high temperature resistant thread, preferably Silicon Carbide thread.

UNITED STATES PRIORITY PRIORITY APPLICATIONS

This application is a continuation-in-part of co-pending ProvisionalApplications Serial No. 60/146,926 High Temperature Capable Ballute ForSpacecraft Deceleration, filed Aug. 3, 1999

BACKGROUND OF THE INVENTION

1. Field of the Invention

The invention relates to the field of spacecraft and, in particular, toan improved ballute for decelerating the spacecraft such that it can beplaced in orbit about an extraterrestrial planet or for actual reentryinto the planet's atmosphere.

2. Description of Related Art

Deceleration of spacecraft for planetary orbit or re-entry into theplanet's atmosphere creates severe temperature and pressure loads on thespacecraft. The space shuttle uses ceramic tiles, which provideexcellent protection. It is especially suited for vehicles that arereusable, even though extensive maintenance procedures are required.Other single reentry spacecraft such as planetary probes havesuccessfully employed heat shields made of ablative materials, whileothers have used metal heat shields. However, such systems impact thesize and weight of a reentry vehicle. All are critical on spacecraft tobe used for interplanetary exploration. Both directly impact size (andtherefore cost) of the launch vehicle required to place a payload intoorbit or to send the spacecraft to another planet.

Inflatable aerodynamic braking devices, commonly called ballutes, whichhave been in existence for a long time, offer a serious alternative.They have the advantage of being storable in a relatively small volumeand, when inflated, can be expanded to a size many times in size. U.S.Pat. No. 4,504,031 “Aerodynamic Braking And Recovery For A SpaceVehicle” and U.S. Pat. No. 4,518,137 “Aerodynamic Braking System For ASpace Vehicle” both by D. G. Andrews disclose an inflatable brakingdevice, which in the stored condition is mounted on the aft end of thespacecraft about a rocket nozzle. When deployed it produced a largeaerodynamic braking surface. The rocket engine exhaust during brakingprovides a cooling layer of gas forward of the braking device such thatit does not overheat. This will allow the spacecraft to enter a lowearth orbit. However, it depends upon the use of a rocket engine toprovide protective cooling gases. It would be unsuitable for a ballistictype reentry into the atmosphere required for landing the spacecraft onthe planet's surface. This is primarily due to the high heat loads thatwould be introduced into the rocket engine as the spacecraft descendedinto the denser atmosphere. In fact, ballutes are generally designed forslowing a spacecraft into orbit about a planet, not for descent to theplanet's surface. Heretofore the prior art ballute designs did notaccommodate such a reentry. However, at least they offer a reduction insize and weight of the spacecraft when used in combination with a moreconventional aerodynamic breaking device such as an ablative heatshield.

High temperature ceramic thermal blanks have proven useful onspacecraft. For example, fibrous silica bafting sandwiched betweensilica fabric and glass fabric. More recently, multi-layer materialshave proven useful. For example, U.S. Pat. No. 5,038,693 “CompositeFlexible Blanket” by D. A. Kourtides, et al. discloses a multi-layerinsulation blanket. It comprises outer layers of Aluminoborosilicate(ABS), with Aluminum and Stainless Steel foils reflective layersthere-be-hind. Spacers made of ABS or polyimide scrim were placedbetween the metal foils. An alternate approach used metal-coatedpolyimide cloth instead of metal foils. The various layers were sewntogether with Silicon Carbide thread. Various combinations and numbersof layers were tested; all of which proved successful. If metal foilswere used, a layer of ABS was inserted therebetween. If metal-coatedpolyimides were used they were aligned such that the metalized sidecontacted the polyimide side of the next layer. However, such insulationwas not designed to be used in an inflatable structure. First of all,there is no specific gasbag layer. Inflatable aerodynamic brakingdevices or ballutes must be internally pressurized during reentry tomaintain the aerodynamic shape. In addition, they must be flexibleenough to store in a compact shape and expand to a comparably largeaerodynamic shape. Also, they must be able to retain their physicalshape and not grossly deform during entry/reentry, which would changeits drag characteristics.

Thus, it is a primary object of the invention to provide an improvedmaterial for an aerodynamic braking device for a spacecraft.

It is another primary object of the invention to provide an improvedmaterial for an inflatable aerodynamic braking device for a spacecraftthat allows the braking device to be stored in a small volume and whichcan be expanded to provide a large aerodynamic braking surface.

SUMMARY OF THE INVENTION

The invention is a material for use in an inflatable aerodynamic brakingdevice for a spacecraft. In detail, the invention includes a pluralityof high temperature resistant barrier outer layers. Preferably, theouter layer is made of Aluminoborosilicate fabric. At least one carboncloth layer is disposed behind said outer layers. A plurality of thermalradiation barrier intermediate layers are disposed behind the at leastone carbon cloth layer. The intermediate layers include a plurality ofmetal foil layers, at least one additional carbon cloth layerinterspersed between the metal foil layers. A plurality of metal-coatedpolymeric layers is disposed behind the metal foil layers, preferably apolyimide fabric coated with Aluminum. The metal foil layers areAluminum and Titanium. At least one inflatable gas barrier interiorlayer is disposed behind the intermediate layers, preferably made of apolyimide material. The outer, intermediate and interior layers are sewntogether by high temperature resistant thread, preferably SiliconCarbide thread.

The novel features which are believed to be characteristic of theinvention, both as to its organization and method of operation, togetherwith further objects and advantages thereof, will be better understoodfrom the following description in connection with the accompanyingdrawings in which the presently preferred embodiment is of the inventionis illustrated by way of example. It is to be expressly understood,however, that the drawings are for purposes of illustration anddescription only and are not intended as a definition of the limits ofthe invention.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of a spacecraft designed for re-entry intoa planetary atmosphere.

FIG. 2 is a perspective view of the spacecraft shown in FIG. 1 with theinflatable aerodynamic breaking device inflated.

FIG. 3 is a cross-sectional view of the spacecraft shown in FIG. 2.

FIG. 4 is an exploded partial cross-sectional view of the breakingdevice shown in FIG. 3 illustrating the construction thereof.

FIG. 5 is a partial view of the device illustrated in FIG. 3.

FIG. 6 is a table summarizing the number and type of layers of apreferred material assembly for the braking device.

FIG. 7 is a graph of the performance of the material assembly shown inFIG. 4.

DESCRIPTION OF THE PREFERRED EMBODIMENT

Referring to FIGS. 1, the spacecraft, generally indicated by numeral 10,includes a payload section 12, heat shield 14 and an inflatableaerodynamic braking device 16 illustrated in the stored or collapsedcondition secured with a releasable band 18. In FIGS. 2 and 3, thebraking device is illustrated in the inflated condition and indicated bynumeral 16′, with the band 18 released therefrom. The retention band 18,for example, could be a simple band secured by explosively actuatedbolts (not shown). The device 16′ forms a smooth extension of the heatshield 14 at the interface 19. The placement during storage, method ofinflation, and retention method are not deemed critical to the inventionand will not be discussed in further detail.

Referring now to FIGS. 4-6, the device 16 is composed of three mainsections: an outer high temperature abrasion resistant portion 20, anintermediate radiation reflection section 21 and a gas barrier section22. Each of these sections 20, 21, and 22 are composed of one or morelayers of material. The outer high temperature abrasion resistantportion 20 typically includes one or two plies of Aluminoborosilicate23A and 23B. For example NEXTEL®, Style AF-12, 7.9 oz/sq.yd., 0.012 inchthick, 1200 denier warp and fill, 5 harness satin weave. Such a materialis available from 3-M Ceramic Fiber Products, St. Paul Minn.Aluminoborsilicate provides excellent abrasion resistance and canwithstand temperatures of around 2000 degrees C for the time periods ofinterest during reentry.

The intermediate radiation reflection section 21 comprises amulti-number of plies. Immediately behind the Aluminoborosilicate pliesis a carbon cloth ply 25, for example, Zoltex Panex 30 SW08, 8 harnessSatin weave, 8.5 oz/sq.yd., 0.040 inch thick. Such material is availablefrom Zoltec Corporation, St. Louis, Mo. The carbon cloth ply 25 acts asa heat sink impeding the heat flow into the remainder of the device.Next is a single ply of commercially pure Titanium metal foil 26, 0.002inch thick, followed by another ply of carbon cloth 28. Thereafter threemore plies of Titanium foil 30, 32 and 34. Preferably, the Titaniumplies 30, 32 and 34 are dimpled so that they are substantially spacedfrom each other. This is necessary because the metal foil plies act asradiation reflectors and to be effective must be spaced from each other.Titanium, while acting as a reflective layer, also has poor heattransfer characteristics, and thus also serves as a heat sink.

Thereafter, four plies 36, 38, 40 and 42 of dimpled Aluminum foil, whichalso act as reflective layers, are disposed behind the Titanium ply 34.The Aluminum foil is preferably 0.005 inch thick, No. 1145 (99.45% pureAluminum). A single ply 44 of Aluminum metallized polyimide, 0.002 inchthick, followed by 11 plies 46, 47, 48, 49, 50, 51, 52, 53, 54, 55, 56,of Aluminum metallized polyimide, 0.003 inch thick, followed a singleply 58 of Aluminum metallized polyimide, 0.002 inch thick. All themetalized plies of polyimide are metal-coated side on both sides.However, single side coating may suffice in many applications. Thepolyimide is preferably manufactured by DuPont High Performance Films,Circleville, Ohio. Finally, the gas barrier layer comprises a single ply60 of polyimide film, 0.005 inch thick, and acts as a pressurizablebladder, again preferably KAPTON®.

The number of plies will very depending upon the severity of thereentry. However, and material assembly will require: the outer hightemperature abrasion resistant portion 20 of one or two plies ofAluminoborosilicate 23A and 23B; the carbon cloth ply 24 therebehind; aplurality of metal foil layers with an additional carbon cloth plybetween the metal layers; a plurality of metallized polyimide layers;and the polyimide bladder 60. For example, in the ballute application,this material would be needed in proximity to the heat shield 14.Further out toward the periphery of the device 16, the heating is lowerand a less complex material assembly is required.

Because the device 16 is made up of a large number of layers (25 in theexample provided); it may be desirable, if not mandatory to stitch thelayers together. As illustrated in FIG. 5, this can be accomplished bysewing the layers together with thread, such as one made of SiliconCarbide. A suitable Silicon Carbide thread is disclosed in thepreviously discussed US Patent U.S. Pat. No. 5,038,693 “CompositeFlexible Blanket”.

Theoretical studies of an entry to the atmosphere of the planet Mars,have indicated that the heating rate on a device 16 for a spacecraft 10would significant. As summarized in the graph presented in FIG. 7, thesubject device 16, while experiencing outer surface temperatures in the1100 degree C range at the device 16′ and heat shield interface 19 witha peak heating rate of 35 w/cm², would be capable of limiting thetemperature at the bladder 60 to less then 200 degrees C.

While the invention has been described with reference to a particularembodiment, it should be understood that the embodiment is merelyillustrative, as there are numerous variations and modifications, whichmay be made by those skilled in the art. Thus, the invention is to beconstrued as being limited only by the spirit and scope of the appendedclaims.

INDUSTRIAL APPLICABILITY

The invention has applicability to the spacecraft manufacturingindustry.

What is claimed is:
 1. A material assembly for use in an inflatableaerodynamic braking device for a spacecraft comprising: a plurality ofhigh temperature resistant barrier outer layers; a least one carboncloth layer disposed behind said outer layers; a plurality of thermalradiation barrier intermediate layers disposed behind said at least onecarbon cloth layer, said intermediate layers comprising: a plurality ofmetal foil layers; at least one additional carbon cloth layerinterspersed between said metal foil layers; a pluralitypolymeric-layers behind said metal foil layers, said polymeric layersmetal coated on at least one side; and at least one inflatable gasbarrier interior layer disposed behind said intermediate layers; andsaid outer, intermediate and interior layers joined together at theperiphery.
 2. The material assembly as set forth in claim 1 wherein saidplurality of high temperature resistant barrier outer layers comprisesaluminoborsilicate fabric.
 3. The material assembly as set forth inclaim 1 wherein said metal foil layers include at least one layer ofAluminum foil.
 4. The material assembly as set forth in claim 3 whereinsaid metal foil layers include at least one layer of Titanium foil. 5.The material assembly as set forth in claim 1 wherein said metal coatedpolymeric layers are made of an Aluminum coated polyimide layer.
 6. Thematerial assembly as set forth in claim 1 wherein said at least oneinflatable gas barrier interior layer is a polyimide fabric.
 7. Thematerial assembly as set forth in claim 1, or 2, or 3, or 4, or 5, or 6,wherein said outer, intermediate and interior layers joined together atthe periphery by a sewn joint.
 8. The material assembly as set forth inclaim 7, wherein said sewn joint is accomplished using a thread made ofSilicon Carbide.